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Why was the short field takeoff procedure changed from 1978 (flaps UP) to 1979 (flaps 10°) on the Cessna 172?
The Cessna 172N was produced for model year 1977 to 1980 (according to wikipedia). The first two years the short field standard procedure (i.e. "the checklist") says to set flaps up (or 0°), but last two years the standard procedure was changed to flaps 10°.
Why the change in the short field procedure? Particularly interested in any differences in aircraft / airframe design which may have led to this change, but whatever the reason I am curious.
See images below for screenshots from all four years' operating manuals, as well as the amplified procedures from 1977.
Additional info:

The 1979 model is also the first year where the first notch (that is, 10°) of flaps were permitted up to 110 KIAS.
1977 and 1978 models do not prohibit 10° flaps on takeoff. The earlier models explain in the operating manual's amplified procedures that 10° flaps results in shorter ground roll, but also shallows the climb angle. Therefore, you will have a longer distance to clear a 50 ft obstacle (see last image). Nonetheless, the standard procedure is flaps 0° for those two years.
The 1979 operator's manual contradicts its own standard procedure in the amplified procedure. The checklist has "Wing flaps -- 10°" but the amplified procedure says "Normal and short field takeoffs are performed with flaps up"...
I'm still confirming this, but I believe every model before 1977 used flaps UP for short field takeoff, and every model after 1980 used flaps 10° for short field. The exception is the retractable gear 172 which use flaps UP for all years (again - still going through all the old POHs I can find online to confirm this).
1977 saw a +10 horsepower gain over the 1976 172M model (thanks JohnK).

1977 POH (pg 4-8 / pdf pg 48)

1978 POH (pg 4-8 / pdf pg 28)

1979 POH (pg 4-8 / pdf pg 66)

1980 POH (pg 4-8 / pdf pg 54)

1977 "Amplified Procedure" (pg 4-14 / pdf pg 54)

Was this change for the new model - i.e. were the manuals for older models amended, or did the change just apply to new models?

This is an interesting question. The only change I can think of that would significantly affect takeoff and climb performance is the leading edge droop cuff added to the wing, in 1973 or 74 IIRC, and the change to 160hp in the N model in '77. The engine adopted in '77 was a famous debacle, with the oddly predictive model suffix "H2AD".

@DJClayworth I'm not sure where to look up amendments to aircraft manuals. The only place I know to look is the 172 TCDS (type certificate data sheet) - and I can't find anything regarding short field takeoffs in there.

@JohnK Regarding the 160 hp engine you mentioned: Could it have taken the engineers 2 years to realize the impact on short field performance? I every model before the 'N' used flaps UP, and every model after the 'N' uses 10° flaps for short field. The exception is the retracts (RGs). The RGs using flaps UP for short field may be even more evidence pointing to the engine being the main factor: the RGs are heavier and therefore the +10 hp may not have been enough to benefit from increased flaps.

You could have something were testing results in an update and change to the original performance data, but it would be strange to limit it to certain years unless there was some other configuration change.  Could the change in flap selector to the preselect do it? I think that was introduced in '79. A change from needing to hold the lever to just flipping it could have changed the approach to takeoff profile. Possibly.

One possibility is that during additional flight testing between 1978 and 1979, test pilots found that they saw a decrease in takeoff run distances or best angle of climb for obstacle clearance using 10° of flap as opposed to with flaps up and it was incorporated into flight manuals for the next iterations of the aircraft.

From doing a bit of digging on this, I think you'll find the answer not in the procedures, but in the performance charts. Compare the speeds and distances over a 50ft obstacle in both models with and without flaps.

I cannot find any indication there is a contradiction in the POH statements regarding the take-off procedure being questioned.

@ThomasPerry in the 1979 PDF, compare the short field procedure on page 4-8 (flaps 10°) to the amplified procedure on page 4-14 (short field takeoffs are performed with flaps up).

@codenoob  Ah, yes.  Ok, I think I see your question.  This is apparent traced back to 1956!  Still checking the details, but this is 'live' in the 172S POH.  Evidently this is guidance to inform choices at the pilot's discretion.  As this is mentioned in an advisory manner, the intent of Cessna may have been to preclude adverse consequences arising from any carelessness, incapacity, or neglect.

I think you indirectly answered your question by including the amplified procedure section from the 1977 manual.  Cessna appears to give conflicting instructions on how to perform a short field takeoff. I suspect that they were getting customer feedback concerning the clarity of operating procedures, and then made the changes for the 1979 manual to reflect a uniform procedure.

Personally, I'm more inclined to believe the extra 10 hp was the factor, and perhaps the test pilots didn't register the impact till '79 as you mentioned in your comment above. But these are all just opinions!

@codenoob perhaps more hp changed the best excess thrust curve to a slightly lower airspeed with flaps.   Keep in mind the prop pitch may also figure significantly as thrust = engine power × efficiency.  (Procedures may vary between  climb prop and cruise prop).  A strong clue is the passage:  advantage of flaps is lost in "high altitude takeoffs in hot weather at maximum weight" -  1977 amplified.

@codenoob There is no evidence that the increase in horsepower had anything to do with the amplified procedure.  There does not appear to be an issue here.  Apparently, Cessna changed nothing in the procedures being documented that are later identified as Amplified Procedures...

@ThomasPerry I think you are misunderstanding my question. In the procedure checklist (not the amplified procedure) the short field was flaps 0° from inception to 1978, and flaps 10° from 1979 to present. Why was the procedure (the checlist - NOT the amplified procedure) changed between those years? The business about the amplified procedure can be ignored - I was just pointing it out in case it was in any way a clue to the answer.

...10 degrees of flaps shortens the ground roll, but shallows the climb angle

This seems to be a mirror image of aircraft gliding performance, where lift to drag ratio is highest when the aircraft is "clean".
Flaps are useful in short/soft field takeoffs because they enhance ground effect.
An expert pilot will takeoff with flaps, raise them as the aircraft accelerates, and climb.
A less experienced pilot may run into trouble transitioning out of ground effect with insufficient airspeed.
But more horsepower is very useful for STOL because it allows higher coefficient of lift even if lift to drag ratio is lower$^1$.
This aircraft, similar in weight to the 172, used 240 hp, slats, and flaps to gain legendary short field performance.
$^1$ excess thrust determines maximum angle of climb

All very interesting, as usual from your answers. But alas, again as usual, it's hard to pinpoint the actual answer to the question posed. "Why was the short field takeoff procedure changed from 1978 (flaps UP) to 1979 (flaps 10°) on the Cessna 172?"

Robert, your answer has no downvotes. I refrained from that to give you time to fix your answer to be an answer. Sidenote: I fly 172's very regularly ;) I'm very interested in the answer to this one, hence my interest.

Thanks for being the first to take a crack at the question, but I was really interested in why the procedure changed in the manual between those specific years. As far as "need to know", well, let's just say I "want to know".

OK, ok.  It would probably be liability issues as more and more people with less experience got into flying.  Did not Cessna halt production for a number of years for the very same reason?

This post does not attempt to answer the stated question.

@RalphJ  because of issues such as weather and weight there may not be one correct answer.  More opinions may be found here

@RobertDiGiovanni Read The Full Question... while "which is better" may not have a single objective answer (as shown in the discussion you linked),"why did Cessna change their manual" reflects the fact that they did make that change. This question asks why that happened, and no amount of "I prefer to..." responses provides that answer, unless it's coming from someone who was involved in the decision. Ideally, an answer will be posted giving insight into Cessna's reasoning at that time.


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Has there ever been a plane with reversed flight controls?
Has there ever been a plane which at least one of the primary flight controls is reversed?
For example, pulling back on the stick will cause a dive, or full throttle is achieved by moving the levers backwards towards the pilot.

Yes but I don't have the references at hand.

Do you mean by accident or by design?

The Wright Flyer had a "normal' (but very sensitive) elevator, and you used your hips for roll.  I wouldn't be surprised if one of the Pre-WW1 planes had reversed controls compared to modern aircraft.

They are not airplanes, but overhead-stick gyrocopters work that way...

It may not be what you're asking, but a US pilot flying a Yak (for example) will have to 'reverse' the rudder input on takeoff: left instead of right rudder, because the prop rotates anticlockwise rather than the usual clockwise rotation in US aircraft.

This is the hell of playing Nintendo video games !!!!!!!!!   What they call "normal" is "inverted" by any sane standard.

Arguably, hang gliders and "trikes"--

Well into WWII, quite a few French aircraft had the throttle reversed- with full throttle to the rear. The French also customized the imported aircraft in WWII to this effect. From the book American Warplanes of WWII- Fighters, Bombers and Liaison by Col. John. D. Current:

... Officially designated Curtiss H75-C1, the aircraft ... had instruments calibrated for the metric system, ... a French style throttle which operated in reverse from U.S. and British aircraft (full throttle was to the rear rather than to the front)...

From joebaugher.com about the same aircraft:

The Hawk 75A-1 was powered by the Pratt & Whitney R-1830-SC-G engine, with an international rating of 900 hp at 12,000 feet and 950 hp for takeoff.  ... A modified seat was fitted to accommodate the French Lemercier back parachute. The throttle operated in the "French fashion", i.e. in the reverse direction to the throttles of British or US aircraft.

Looks like the French modified the T-28 Trojans too (though I'm not absolutely sure about this):

... the French were looking for a replacement for their ground attack T-6s. ... The French did however acquire 148 A models that were in storage at Davis-Montham. They then changed the engines, (removed the R-1300 and installed the R-1820-76A), reversed the throttle operation, (forward is closed aft is open ?? why do this?) ... These planes were then used by the French in Algeria from the end of 1960 until the end of the war in July of 1962.

Not only the French, the Italians and Japanese also used reversed throttles, which caused quite a few problems in service, as the Wikipedia page on Macchi_C.205 notes:

In the brief German use, Veltros had at least five losses by accidents, often caused by the inverted throttle used on Italian aircraft (In German and Allied fighters the "open throttle" position was forward, not back, and this was the source of several errors).

The Italians too faced problems with this difference in controls. According to Oddone Colonna, an Italian pilot in 150 Gruppo CT at Gruppo Caccia:

In the summer of 1943 at Vicenza airport I made my first acquaintance with the "Gustavo", a Messerschmitt fighter of the G-6 version. I was among pilots from 23° and 150° Gruppo Caccia to get first flights in the German plane that had been entering service at the Regia Aeronautica for few months.
We had two main difficulties on test taxing at the Vicentine airport.
1)    The gas throttle had an inverted movement compared to Italian vehicles. [The Italians like the French used a reversed throttle in their planes. TNi]
2)    The Plane had hydraulic pedal brakes instead of pneumatic ones with differential command like on other Italian vehicles.

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What is the temperature of the brakes after a typical landing?
On an airliner, what is the temperature of the brakes after a typical landing, and how long does it take to cool down?
Is it hot enough that mechanics should wait before servicing an aircraft that has just touched down minutes before?

You definitely don't want to touch them after a few minutes, you wouldn't want to touch your car breaks a few minutes after heavy breaking.

Most airlines and some fighter jets like the eurofighter have little fans that start to spin on the ground and cool down the brakes. A A320 wheel: http://previews.123rf.com/images/silverfoxz/silverfoxz1507/silverfoxz150700009/42240568-Thailand-June-2015-Airbus-A320-Aircraft-Wheel-at-BKK-Airport-Soft-focus-Stock-Photo.jpg Eurofighter wheel: http://www.aircraftresourcecenter.com/awa01/701-800/awa719-Eurofighter-Watson/18.jpg

@Notts90 even on a pushbike

@jklingler this devices are quite spread among aircraft designed for tiny turnaround times. I read somewhere the temperature must be low enough to handle a rejected take-off, and those devices allow the temperature to be fine in few minutes.

"the temperature must be low enough to handle a rejected take-off", not too high would be more accurate (the brake disks are glowing red). I think the RTO certification test requires fire not extending outside the wheel before 5 minutes after a RTO at maximum weight and minimum runway length (time for firefighters to reach the aircraft).

Not hot enough to melt steel.

Judging by this IATA document, and looking at figure 4 in section 4.2.1, the brake temperature easily exceeds 700°C during landing. They also show some brake temperature history timeline for an airplane without fan cooling.

(Figure 4)

Figure 2. made me laugh.

Welcome to aviation.SE!

In figure 4 they nicely distinguish disk temperature from indicated temperature. Shame that they didn't label the time axis properly. But the numbers could be minutes. It shows nicely that this brakes never get cooler than 100 °C. Even when parked, the exponential heat decay would take hours before dropping below a touch friendly temperature (like 60 °C).

"Do not distribute this outside Airbus" is what I chuckled about.

@Pondlife: Thanks!

@ymb1: Thank you for the great edit -- In general, I recon the axis is also minutes, looking at the taxi in and out times. All in all, I would definitely not touch the brakes!

Great graph! Some suggestions for improvement: Was this measured in the lab, or at an actual aircraft? Which one? You should point out that 700 dC refers to the brake disk, not to the brake unit. What are the actual answers to the original question?

Managing Uneven Brake Temperatures on Twin-Aisle Airplanes During Short Flights was published by Boeing in 2001. The article details the threshold temperatures for the brake overheat warning of  a few different twin-aisle plane types. Assuming that during a typical landing the brake overheat warning is not triggered and the brake system is not excessively over-designed, I would expect typical temperatures around 300 to 400 °C, very much depending on the plane.
The article also states that when leaving the gate, the brakes will still be warm from the previous landing, and that brake temperature may build up during the day. Note that (at least for the brakes shown) brake disks are not as exposed as the ones usually found in cars. Cooling through air flow seems quite limited.
Judging by that, I would say that yes, mechanics should be very careful before touching the brakes of a plane that has just landed. With the wheels mounted, we are talking more about hours, not minutes.

Steel brakes and carbon brakes have substantially different temperature ranges, and wear patterns.
Wear in steel brakes is accelerated substantially at higher temperatures, so (especially when taxiing) recommended brake usage is to use frequent, light applications of the brakes to maintain as close as possible to the preferred taxi speed. Most wear on steel brakes happens during landing, when you have little choice but to apply the brakes continuously, even though the temperature gets quite high.
With carbon brakes, rather the opposite is true: wear depends primarily on the number of times the brakes are applied, not the temperature reached. As such, the normal recommendation during taxiing is to let the speed run up to the maximum allowed, then apply the brakes fairly firmly and continuously once until the speed drops to a specified minimum. With carbon brakes, most wear (~75%) happens during taxiing. As already noted, wear depends primarily on the number of applications, not the temperature, and landing is typically just one application, even though it's a relatively protracted one that produces relatively high temperature.
For example, Boeing recommends as follows:

Because the wear mechanisms are different between carbon and steel brakes, different taxi braking techniques are recommended for carbon brakes in order to maximize brake life.
Steel brake wear is directly proportional to the kinetic energy absorbed by the brakes. Maximum steel brake life can be achieved during taxi by using a large number of small, light brake applications, allowing some time for brake cooling between applications. High airplane gross weights and high brake application speeds tend to reduce steel brake life because they require the brakes to absorb a large amount of kinetic energy.
Carbon brake wear is primarily dependent on the total number of brake applications — one firm brake application causes less wear than several light applications. Maximum carbon brake life can be achieved during taxi by using a small number of long, moderately firm brake applications instead of numerous light brake applications. This can be achieved by allowing taxi speed to increase from below target speed to above target speed, then using a single firm brake application to reduce speed below the target and repeating if required, rather than maintaining a constant taxi speed using numerous brake applications. Carbon brake wear is much less sensitive to airplane weight and speed than steel brake wear.

As such, it's perfectly normal for carbon brakes to operate at considerably higher temperatures than steel brakes. For carbon, typical operating range is about 600-800C. For Steel brakes, the preferred range is more like 400C. In both cases it also depends on what you measure--the temperature shown on the cockpit indicator is substantially below the peak temperature (at least in all cases of which I'm aware).
For what it's worth, steel brakes also cool faster than carbon brakes as a rule (though I haven't been able to find anything solid about exactly how much faster).
Here's a graph from Boeing showing the temperatures of the carbon brakes in a 777:

So yes, even an hour and a half after landing, the brakes are still hot enough that skin contact will result in a blister essentially immediately. As to waiting before servicing, that would depend on what servicing you're talking about doing. Refueling shortly after landing is common, and the brakes rarely affect its safety to any significant degree. Anything that involved removing the main landing gear wheels so the brakes were exposed...probably needs to wait quite a while longer. Other work near the brakes (e.g., servicing the struts) probably merits a little extra care right after landing, but there aren't too many cases where the brakes are likely to lead to a significant delay.
References:

Advanced Ceramic Technologies & Products, The Ceramic Society of Japan.
Operational Advantages of Carbon Brakes, Boeing.
Brake Energy Considerations in Flight Operations, Rob Root (Boeing).

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What would happen if the flight engineer is unavailable on a 747?
Consider a 747-100/200 with a 3 person crew: two pilots and one flight engineer. During the flight, the flight engineer suffers a medical problem and is unable to serve duty for the rest of the flight.
What difficulties would the pilots face for the remainder of the flight, including navigation and landing, with no one attending the flight engineer station in the cockpit?
For example, would one of the pilots get up and sit at the FE station, leaving the other pilot the only person at the controls? Or would the person get up, go to the FE station to read or adjust something, then sit back at the pilot's seat? Or do they not have to move at all?

If the flight is long enough, alternative crew is needed.

@TrebiaProject. that is why I stated "3 person crew" at the beginning of the question.

"The F/E had the fish, now what?" XD

This question has been asked and answered in a slightly different form at How able were pilots of three-person cockpits at running the normal FE work?, and I refer you to that for basic info. Concerning the detail in your question:

What difficulties would the pilots face for the remainder of the flight, including navigation and landing, with no one attending the flight engineer station in the cockpit?

It would be a nuisance for sure, but as long as there were no abnormal or emergency situations (other than the FE's personal situation), a nuisance would be all that it was. However, one could dream up scenarios that could get really sticky without the specialized knowledge that the FE would be expected to have that the pilots might not have (or had forgotten).

For example, would one of the pilots get up and sit at the FE station, leaving the other pilot the only person at the controls? Or would the person get up, go to the FE station to read or adjust something, then sit back at the pilot's seat? Or do they not have to move at all?

The Captain can see the FE panel. Some of the gauges along the aft portion would be a little hard to read. For example, if he wanted to see what the exact pressure of the Potable water was rather than just that there was pressure, he might have to get up to get close enough. Or if a generator fault light came on, he'd probably have to get up to see what the fault was.
There wouldn't be any need to stay in the FE seat. If the flight were long, there would be some trips to the FE seat for fuel management. You can have as many as 9 fuel tanks (7 is the usual), and there are rules for the fuel burn schedule. The most important and most time consuming FE task during cruise was keeping track of what the fuel burn was, especially on the 747-100 because it didn't have the fuel capacity of the -200. For example, a -100 going from JFK to Tel Aviv would sometimes not have the fuel to make it. Abeam Athens the FE would inform the Captain either than they had the fuel to continue or that they would need to divert to Athens for fuel.
If the Captain wanted to talk to company for some reason, it was usually the FE that established contact via HF, and it was usually the FE that answered SELCALs. But the pilots sometimes did it.
Flight attendants might show up and ask for a general cabin temperature change.
Of course, it the FE had a severe medical problem, you would probably land at the nearest suitable airport just as you would for a passenger medical emergency.

@BaileyS yup, that was intended (-;


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Nose gear weights
I’ve been reading about aircraft tugs and how some of them utilise a cradle design to hold the nose/tailwheel so the aircraft can be manoeuvred.
Given my background in electrical engineering, I have to know more as it’s quite an interesting concept. However, I lack much of an aviation background (my only relationship with planes is using them to go abroad) so I’m not familiar with the aviation-concepts revolving weight.
Having looked at the Wikipedia pages of aircraft like the Cessna 152, Piper PA-28 Cherokee and Piper M350, the highest weight (MTOW) is between 1000 and 2000kg. My question is, how much of that weight goes through the nose wheel, and thus the tug?

Does this answer your question?

If you do a really bad job, the weight on the nose wheel will be negative: https://aviation.stackexchange.com/questions/24569/which-aircraft-models-need-a-tripod-jack-for-tail-support-at-airport

A typical small plane tug looks like this: https://upload.wikimedia.org/wikipedia/commons/d/da/Aerotechnik.ev97a.eurostar.g-ccmp.arp.jpg

In a tricycle gear aircraft, the nosewheel supports little of the weight. The center of gravity of the aircraft is only slightly in front of the main gear, and so the nose gear only has to hold the small amount of weight required to keep the aircraft from tipping forward onto its nose.
Exactly how much weight the nosewheel has to support depends on how the aircraft is loaded. It is greatest when loaded to max gross weight, with the center of gravity at its most forward allowable.
As an example, a 1978 Cessna 152 has a max gross of 1670 pounds and a forward center of gravity limit at 32.65 inches* aft of the "datum," an imaginary plane passing through the front face of the firewall. The main wheels are located 47.1 inches aft of datum and the nose wheel is located 10.8 inches forward of datum.
Balancing out the forces and torques for this loading reveals that the main gear needs to support 1253 pounds and the nose gear supports only 417 pounds:
$ 417 + 1253 = 1670 $
$ 417(10.8)+1670(32.65)=1253(47.1) $
It tends to be even less for a tailwheel aircraft. Since the tailwheel is so far from the center of gravity, it has a long lever arm and very little force is required to produce a large torque.
*Note: at lighter weights, the center of gravity can be more forward than this, up to 31 inches aft of datum at 1350 pounds. At these loadings, the nose gear will support a higher percentage of the overall weight, but since the overall weight is smaller, the total weight supported by the nose gear is smaller. Example: at 1350 pounds with a center of gravity 31 inches aft of datum, the nose wheel supports 375 pounds. At 28% of the gross weight, this is a slightly larger percentage, but it is still a smaller number. It's possible that some aircraft's weight and balance requirements are such that the maximum loading on the nose gear is at some weight other than max gross.

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How much force do you need to tow a P68 Vulcanair?
I found an old thread discussing the force required to pull an airliner, which was far less than I thought it would be. I wanted to compare this to a smaller aircraft (approximately 2000kg) but couldn't find anything. Do any of you know what the answer to this would be?
Edit: if anyone can provide the answer in hp or in lb/f or in Nm, that would be greatly appreciated as that’s what the older thread uses (makes the most sense to compare directly in the same units) Thanks!

For a 3000 pound airplane, about 700 pounds, but likely much less.
A handheld tug
for a similar Beechcraft Baron B55 (twin engine, tricycle gear, 3263 pounds, 6 seats)
uses a cordless drill to apply
1150 inch-pounds of torque to a 3 1/4 inch diameter drive roller.
That works out to a force of 708 pounds.
Online specifications for its DeWalt 20-volt brushless drill are vague,
but if we assume a 1 hour run time from its 8 amp-hour battery,
that's 8 amps times 20 volts = 160 watts of power,
not too much for one human to exert.
However, this tug manufacturer offers it to replace

busting your back moving your bird in and out of your hangar by hand

and NASA reports that a human's maximum horizontal pushing force is about 1500 N or 340 pounds.
The Canadian Centre for Occupational Health and Safety advises that in the workplace, horizontal pushing or pulling should not exceed 50 pounds (165 pounds when "a worker can support their body (or feet) against a firm structure").
So if humans regularly pull such airplanes unassisted, the tug is overengineered, and the actual force needed may be well under 200 pounds.


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How thick is the skin of an aircraft like the Airbus A350 or Boeing B777?
We can evaluate how thick walls are by looking at door or window frames, however that isn't representative of an aircraft fuselage, which appears to be very thin:

Airbus A350 XWB forward fuselage, source

How thick (thin!) is the fuselage?
Do some areas have a thicker skin for some reason?
What about glass, doors, wings, floors, and other elements than the walls?

(I'm talking about the aircraft skin itself, not stringers or other structural stiffeners.)

It is really quite thin and fragile too. You see the red robotic arms? They are equipped with load sensors and once the load is exceeded they go into floating mode with an alarm sounding and those green lights that you see everywhere go into orange or red failure mode. It is a rather bad idea to temporarily disable these sensors when testing the software to compute the robot arm movements and mess up and then you can easily distort a part of the fuselage (insurances can be a nice feature)

You can't evaluate the thickness of a building's walls by examining the door or window openings. Like aircraft walls, building walls are composite structures. Masonry walls have inner and outer leaves with an insulated void, timber-framed walls often have an outer shell of fairly thin siding. You are asking about the thickness of those outer skins, not of the structural elements or the whole.

Imagine an aluminum pop can - blow it up to the size of a 737 fuselage and the aircraft skin will be about half to a quarter of the thickness of the pop can wall thickness.

@mins  No, that's not even close.  Maybe I wasn't clear.  If you scaled a standard 355mL beverage can proportionally to a cylinder of diameter equal to the fuselage of something like a B737 (imagine exactly a pop can the size of an aircraft), you would find that the pop can would have a wall thickness several times (2-4) thicker than the aircraft.  This is just to give a sense of scale with respect to something you might handle in your everyday life.

@PlasmaHH: Interesting information. Having a second look at the image, I realize there are plenty of questions that could be asked about the way a section like this is built without damaging the panels.

@mins If I had time, I could probably answer a lot of them, at least for Airbus A340/A380/A400M, but you know, life and kids and stuff...

Below is a cross section of the skin around the passenger windows of a Boeing 747-100 (N602US). It's 2 layers, each layer is ~.075" (1.9 mm) thick. This is the first iteration of the 747 (50 years ago) so might be different on later series as they improved the design and reduced weight.

Source: own work

Welcome to Av.SE!

What a surprise when they wanted to takeoff, and discovered a pax from the previous flight had managed to cut out 1 m² of the skin... +1

Welcome indeed! One thing asked 'round these parts is that you credit your photos. Please indicate that it's your pic, or include a link to where you found it. Take the [tour] and browse the [help] to get a better understanding of what's expected.

@FreeMan couldn't find an example of self attribution but thought I saw it worded this way before. Let me know if thats not right. Thx!

That works well enough for me!  Nice pic, BTW.

@FreeMan thx! :D

From whats written here

Fokker F100 skin is 1.0 - 1.4mm
Boeing 747 skin is 1.8 - 2.2mm
Airbus A320 skin is about 1.1mm

According this article on the Boeing 757

...Although Boeing specified that the skin in that area of the
fuselage must be 0.039in (0.99mm) thick...

Some interesting info here as well.

Have any figures for composite fuselages?

@NickT I have come across a few things on that topic. I think that might be a great question all to its self as composite skins are much different than metal skinned aircraft.

The first source (forum) also states: "More recent commercial a/c have fuselage skins that are quite a bit thinner than 1/4 inch". I would prefer figures to be supported by technical documents (if possible). The A350 fuselage is 50% composites.

Just a couple of mm thick, about 2-3 is enough. The fuselage barrel gets its strength from the large diameter: bending moments of the fuselage (and wing) get translated into tension and compression loads in the skin. The aircraft skin is the main load bearing structure.

It's the same principle as that for I-beams:they are that shape because the top and bottom plates carry the useful load, the middle bit is basically mainly there to keep them apart. If a beam is required for more load, increasing the distance between top and bottom has a quadratic effect on load level reduction. And the same principle that nature applies to bird bones: they are hollow, because the outer bit is the most effective.
The stringers and other structural helpers are mainly there to help keep the skin in shape, prevent buckling etc. Some older and/or smaller wings use a single, I-beam type spar for absorbing the bending moment, but most effective is the wing box shape: front spar, rear spar, top skin, bottom skin. It resists torsion best and reduces flutter, it also keeps the skin in better shape.

The fuselage skin varies in thickness depending on the section, physical features, structural loading, etc. but is between about 2-4mm in thickness for pressurized aircraft and roughly half that value for unpressurized airplanes.

Actually it even varies along a single panel. In this movie (from a german children television show) it can be seen that the panels of an A320 are actually etched to different thickness depending on the proximity to stringers or windows: https://youtu.be/M83SZfATzfs?t=251

Yes areas of the skin can have doublers in sections determined to have high stress concentrations such as around windows, doors, spars, frames, etc.  The thickness of the skin as well as all other structural members is determined by iterative stress analysis during design.

Well I don't would call this doublers, as actually there is nothing joined or attached, just the whole skin is etched down to three or more different thickness levels depending on the location. So the answer is much more complicated than one thickness value...

No that’s the industry term for it.

The skin on the wings also tends to be a bit thicker, as well as being composed of different materials and finishes.
MD/DC, Airbus, and Boeing have aluminum skin roughly an 1/8th inch thick that is backed with a sealed fiberglass coating which provides the skin with strength and flexibility. The underbellies have slightly thicker skin and so do areas subjected to high wind.
Areas like the APU intake and gear doors have significantly thicker skin up to a 1/4 inch.

That image is misleading. It does not show a solid slab, but a hollow box section.

Do you have any links for further reading?

I have two keychains both with pieces of fuselage from Virgin Atlantic 747-400's. One is 1.6mm thick and the other is 2.8mm thick!

At minimum most external aircraft skins I've every worked with are 14swg (2mm) thickness having had to wheel & structurally fit these personally.
At no time has there ever been an outside aircraft skin that has ever been below 16swg (1.66mm) as it structurally just would not be enough.
Someone stated 0.039" (20swg/1mm) as an outside skin when usually this often is for the inner layer skin where an outer layer is also fitted.
A 20swg/1mm skin would increase fatigue or corrosion problems vastly & given the altitudes large jets such as the Boeing 747 fly at this would itself be a disaster waiting to happen & a design failure.


text:
Why is my WAAS equipped aircraft telling me to descend sooner than the FAF (KDVT 25L LPV)
I ran into a weird issue today with our G1000 WAAS-equipped aircraft. The aircraft wanted me to begin decent .9nm before the FAF on the KDVT 25L LPV approach. The LPV is considered a non-precision approach (Edit: it is not an NP approach, it is a APV which would have you follow GS when intercepted at 4700), so you should fly at 4700 feet to FEGEV and then go capture the glide slope, so this has the potential to screw up a student or anyone for that matter by following the glide slope almost a mile early and getting a potential Pilot Deviation.
I asked ATC what they showed our altitude at and was correct all three times.
I could not find the proper FAA source quickly, but Honeywell had this to say:

An LPV glidepath is calculated entirely using SBAS altitude. That is, the GPS constructs the glidepath using GPS altitude and height above earth ellipsoid to produce a path that unlike baro-VNAV, is not susceptible to pressure and temperature variations. This means that like an ILS, an LPV glidepath is essentially a fixed path in space being tracked using GPS altitude to a DA that the pilot reads from a baro-altimeter. This is a component to how an LPV can achieve minimums comparable to an ILS.

So if they are both fixed paths in space, why is one of them having me descend so early?
I didn't know if anyone else had run into this issue before. Is it due to a faulty GPS receiver? (both GPS 1 and 2 were showing the same)
Any insights are appreciated.
(Edit: Concern is how early it is having us capture the GS where someone on the LNAV for the same approach would be at 4700 feet for almost a mile after we begin descending on the LPV)
On ground GPS signal and altimeter altitude match:

KDVT 25L LPV approach on glide .9nm ahead of FAF:

KCGZ ILS is right on (.1nm off) at correct altitude for FAF:

Approach plate:

Looks like Chris added the approach plate. Thank you, Chris, and happy to be here.  Seems like a great place to learn.

Was this approach flown recently and in the afternoon?  Given that KDVT temps are currently 20-35°F above "typical", that means when flying at 4700 ft baro. alt. you're 100+ feet above "typical" when measured geometrically.  Wouldn't that push back how soon you intercept the glideslope, which doesn't change with temp, by at least a quarter nautical mile?

@CodyP, Yes it seems that from what Chris calculated the equipment is actually working perfectly and that since it was approx ISA plus 20-25 yesterday at that altitude the GPS altitude was almost 300 feet different than our indicated (Static) in the aircraft causing us to intercept the GS .9nm early. It's one thing to learn that going through training and then seeing it in practice. Was shocked by how "far" off it was from the standard LNAV FAF point.

That's what being 27 degrees above ISA will do! For a quicker estimate of how much effect the tempertaute has- ISA temperature is about 285K at TDZE and the actual temperature was around 312K, about 9% higher. So it makes the glideslope intercept about 9% farther. FEGEV is 9.7 NM away from the threshold, so that 9% is about .9nm. Roughly every 3 degrees from ISA changes things by 1%.

RE: your concern added at the end. This is an advantage of the LPV. Flying the LNAV on a hot day requires the pilot to fly a steeper descent than they otherwise would have to in order to avoid busting the altitude restrictions at FEGEV and GIGEC. The LNAV pilot will fly a higher (and thus steeper) approach as the temperature rises, while the LPV pilot can fly basically the same approach regardless of temperature (aside from the DA, which will be higher on a hotter day).

The G1000 is correct. Just like you would on an ILS, once you have intercepted the glidepath you should follow it down. The altitude restriction at FEGEV (the final approach fix) and the stepdown altitude at GIGEC apply to the LNAV version of the approach, but not the LPV.
The AIM touches on this in 5-4-5(b)(5):

Altitude restrictions depicted at stepdown fixes within the final approach segment are applicable only when flying a Non-Precision Approach to a straight-in or circling line of minima identified as an MDA (H). These altitude restrictions may be annotated with a note “LOC only” or “LNAV only.” Stepdown fix altitude restrictions within the final approach segment do not apply to pilots using Precision Approach (ILS) or Approach with Vertical Guidance (LPV, LNAV/VNAV) lines of minima identified as a DA(H), since obstacle clearance on these approaches is based on the aircraft following the applicable vertical guidance. Pilots are responsible for adherence to stepdown fix altitude restrictions when outside the final approach segment (i.e., initial or intermediate segment), regardless of which type of procedure the pilot is flying. (See FIG 5-4-1.)

This mentions that you can ignore altitude restrictions on the final approach segment. So the question then becomes- is FEGEV part of the final approach segment, or not?
The answer is that the final approach segment of an LPV begins as soon as you have intercepted the glidepath, just like an ILS. From TERPS 2-6-1:

The [final approach] segment begins at the PFAF and ends at the MAP
and/or DA.

The PFAF is the "precise final approach fix," defined as:

a calculated WGS84 geographic position located on the final approach course where the designed vertical path (NPA procedures) or glidepath (APV and PA procedures) intercepts the intermediate segment altitude (glidepath intercept altitude). The PFAF marks the beginning of the FAS. The calculation of the distance from LTP to PFAF includes the earth curvature.

The LPV is an APV (approach with vertical guidance). So if you are at the glidepath intercept altitude (4700 feet) and you intercept the glidepath, then you have identified the PFAF. This means you have started the final approach segment and can start your descent along the glidepath, ignoring any further altitude restrictions.
You are still required to adhere to the altitude restrictions at CIPLU and DUYAC, as these are not on the final approach segment.
On hot days, your altimeter reads lower than your true altitude. Thus you are farther from the ground that your altimeter would lead you to think, and you will intercept the glidepath farther out than you would on a standard day. The size of this error can be determined using this lovely formula:
$$ H\times\frac{15-(T+L\times E)}{273.15 + (T+L\times E) - 0.5\times L\times (H+E)} $$
where $H$ is your height above the aerodrome, $E$ is the aerodrome elevation, $T$ is the aerodrome temperature in degrees Celsius, and $L$ is the lapse rate. Plugging in 1475 feet for the aerodrome elevation, 3225 feet for the glidepath intercept height, 39 degrees celsius from yesterday's METAR for KDVT, and assuming a standard lapse rate of $1.98$ degrees per 1000 feet, this gives about 280 feet, so your true altitude MSL when your altimeter reads 4700 feet is actually 4980 feet.
With a glidepath angle of $3^\circ$, this difference in altitude gives you an expected glidepath intercept $280~\text{feet}/\tan(3^\circ)\approx .88~\rm nm$ outside FEGEV. So your equipment appears to be working perfectly!

I appreciate the insight on the TERPS. I missed that in there and got lost in the fact that the LPV is APV (not a Non-PA as I mentioned) and you would follow the GS. It can not be a PA since it is created onboard the aircraft, not from an external source. My concern is that it should be closer than .9nm though correct? Unfortunately, the other aircraft we have that has WAAS is in MX so can't go up and see what that one does with the same day/conditions currently.

@JonathanWilkinson The glidepath angle is 3 degrees. .9nm times tan 3° gives 287 feet error at 3225 feet AGL. Yesterday's metar gives KDVT's temperature as 39 degrees C, which gives a temperature error of about 280 feet. So that's right on the nose! You can expect large errors like this from airports in the summer in Arizona.

@Chris Thank you so much for your help. Out of curiosity, is there any reference you know that would point out the temperature error of the LPV system (only the intercept point) that I could reference/add in the answer above from the FAA, or is the formula sufficient? I know of the snowflake on charts for cold temperature airports but didn't even think of AZ +25 ISA at times causing similar issues. Again, thank you for your help.

@JonathanWilkinson No problem! People don't worry about high temperatures causing issues because they mean you are farther from the ground than expected, so they just give you "bonus" obstacle clearance. The formula is the same for both but the FAA doesn't bother allowing you to lower MDA/DAs on hot days.

@Chris - cool! Something about "5 out of 3 pilots getting things backwards" comes to mind! Since you've edited your post removing the swap, I've deleted the other comments addressing the (now removed) confusion. Thanks!


text:
What is monotonous and non-monotonous flow?
I'm not an engineering graduate so I'm learning as I go along. I've just read a sentence that states:

The transition shapes are identical, while the variety of growth rates of intermittency
are non-monotonous with the increase in aerodisk length

My exhaustive google search does not show any results that in a basic form explain what is "non-monotonous" as well as how its related to intermittency
Can someone please explain what this is?
I read this from here: https://www.mdpi.com/2226-4310/9/12/742

Hi! I think a bit more background could get you a better answer. For starters, where did you read this?

Updated description showing reading material with the above quote

I suppose they meant monotonic

There is no such thing.
In this paper, wherever the adjective monotonous appears, it modifies growth rate, not flow.
Also, the word should be monotonic, not monotonous.
Although MDPI is no longer regarded as a predatory journal, it is still not respected.  A gaffe like this shows that its reviewers are shoddy.

Thanks for this, very helpful. A sad waste of time on my behalf, weeks of breaking down learnings wasted on such a journal :(


text:
How can the derivative of the geodetic angles (lat, lon) be expressed by the Velocities in the NED-Frame?
I have found in the book "Unmanned Rotorcraft Systems" following expressions (eqn. 2.25 and 2.26).

https://www.newbooks-services.de/MediaFiles/Texts/4/9780857296344_Excerpt_002.pdf
$\lambda$ = geodetic longitude
$\phi$ = geodetic latitude
But I don't understand how this expression can be derived or where it originates from. In the book they only say it can be derived by using spherical triangles, but do not show how.
I have found a similar equation in the derivation of the transport rate between ECEF-Frame and NED-Frame in this paper:

But unfortunately it is also not explained, how he got to after the second "="
https://www.researchgate.net/publication/254001148_New_Calibration_and_Computing_Method_for_Direct_Georeferencing_of_Image_and_Scanner_Data_Using_the_Position_and_Angular_Data_of_an_Hybrid_Inertial_Navigation_System

Would you add a comment about what $N_E$ and $M_E$ stands for? But I suspect this is simple trogonometry...

$N_E $ is the radius of the curvature in the prime vertical of the meridional plane.
$ M_E $ is the radius of the curvature in the meridional plane.
See also the first link , p.25 eqn 2.2 - 2.7
(https://www.newbooks-services.de/MediaFiles/Texts/4/9780857296344_Excerpt_002.pdf)

The relationship between peripheral speed and angular velocity in 3d is given by
$\vec{v}=\vec{\omega} \times\vec{r}$
therefore in the 2D case this is
$v = r \cdot \dot{\alpha}$
with $v$ being the velocity, $r$ the distance between the center of rotation and the object and $\dot{\alpha}$ being the angular velocity. Solving this for the angular speed we arrive at:
$\dot{\alpha} = \frac{v}{r}$
Given these relationships, we can easily see that in the velocity north case, the angular velocity is simply $\dot{\phi}$, the velocity $v_{nv}$ and the radius the (mean) radius of the earth plus the flight height $r=N_E+h$. Inserting this into the equation above exactly yields the formula you asked about.
For the east case, a $cos(\phi)$ term is inserted into the sum for the radius (resulting in $r=(M_E+h)\cdot cos(\phi)$, as the radius decreases with increasing latitude.

@Christian Mändel I hope this answers your question sufficiently well. If not, I would have to draw a sketch, but it should be obvious that this is simply the equations for peripheral speed ("Umfangsgeschwindigkeit" in German) applied to the earth.

Thank you ! I was having trouble imagining it. You made it very clear. Cheers!

@ChristianMändle thanks. If that explains the question, please upvote and accept...


text:
Are controllers aware of TCAS RA taking place?
When the TCAS issues a Resolution Advisory (RA) which contradicts an ATC instruction, the RA takes priority.

Are controllers aware (notified on their screen) that a RA has been issued between two converging traffic?
If yes, are the controllers also able to see the RA instruction issued to each specific aircraft?

Is there a difference between TCAS and any other traffic system including looking out the window?

Exactly because of this the Überlingen mid-air collision took place. The TCAS issued an RA to the crews, the ATC stepped in and gave exactly the opposite orders. One air crew followed TCAS, the other the ATC - both planes climbed and collided, killing 71 people.

ATC was and is still generally not aware of RAs before being informed by the pilot
In general ATC is not aware of any RA occurring in aircraft, and is neither aware of the RA content. The crew informs ATC using a phraseology described in ICAO Doc 4444 PANS-ATM, Chapter 12, Para 12.3.1.2:

When the RA occurs, and the crew must deviate from ATC instructions, they inform ATC with "TCAS RA" when possible.
If an ATC instruction is received, which contradicts a current RA, they inform ATC with "Unable, TCAS RA".
The crew then informs ATC when the RA has ended, with "Clear of conflict, returning to (the ATC clearance)", and "resumed" when done.

See instructions to ATC and RTF Phraseology associated with TCAS RAs.
Pilot and ATCO obligations, impact of RAs on responsibilities
RA always takes precedence on ATC and must be followed by the crew (exceptions are investigated). This is because even if the crew has the intruder in sight and another safe maneuver might be executed as well, there is no certainty the visible aircraft is the conflicting one, or is the only conflicting one.
When informed, ATC doesn't provide further instructions to this aircraft before receiving the "clear of conflict" confirmation.
Two issues should not be overlooked:

According to existing regulations, a controller is not relieved from providing a guidance to pilots until informed by the pilot of a RA conflict ("unable TCAS RA" -- a pilot doesn't have to report RAs not preventing to comply with received ATC clearances). Therefore the (legal) responsibilities and obligations of the controller are unclear in the period between the display of the down-linked RA, which can be delayed by technical limits, and the actual report by the pilot which is not instantaneous.

Not all RAs are relevant to ATC, only those which pertain to collision avoidance or conflict with ATC clearances, therefore some filtering must be done either prior to download, or after reception to eliminate "Monitor Vertical Speed", "Maintain Vertical Speed", etc. This filtering may be imperfect and can lead to RA not displayed.

Last but not least, the current view of the controllers federation:

IFATCA is opposed to down linking of any advisories generated by ACAS.
If down linking of ACAS Resolution Advisories becomes mandated, then
IFATCA can only accept this provided that the following criteria are
met:

Clear and unambiguous controller legal responsibilities;
Downlink without delay;
ATC system to be able to receive, process and display the down link to the appropriate control positions;
Compatibility with all ground based safety nets;
Nuisance and false alerts must be kept to an absolute minimum;
ACAS should only be considered as a ‘safety net’.

RA downlink decisions and implementation progress
Due to the responsibility and limitations issues, the adoption of RA display at controller working position is variable among countries, in spite of technical studies confirming the interest (detailed further down), and modern ATC systems being able to receive and manage downlinked RA. Current status for some countries (source):

USA and UK and France don't use it (UK CAA clearly forbids any use).
Czech Republic and Hungary have tested the technology early and use it.
Luxembourg is conducting studies for using it.

Where it is used, the RA data is only for information, and doesn't change ATCO responsibility for collision avoidance, e.g. for Hungary the regulation has been complemented (source):

The TCAS RA Downlink shall be ignored from an operative point of view.
In  other words, it is only for situational awareness and informative
purposes, and the controller shall not consider RA Downlink when
issuing a new  clearance.

UK CAA forbids ANSP to use RA downlink (ibid):

There are currently no ICAO, European or UK pilot procedures, ATC
procedures or legal responsibilities for the use of ACAS RA Downlink.
However, this subject is under consideration by Eurocontrol.
There is no assurance as to the integrity of the RA downlink and
absent or false downlink data could be a possibility.
A downlinked RA without adequate discrimination may lead the ATCO to
inappropriately cease the provision of ATC instructions. However,
under current ICAO procedures, the controller will continue to provide
ATC instructions during RA events that do not deviate from the
clearance or instruction unless such a clearance is at variance with
the RA, at which point the pilot will report ‘UNABLE, TCAS RA’.

Studies conducted to include ACAS RAs in ground-based safety nets
There has been several accidents involving RAs not followed by crews, among them a tragical collision over Switzerland (2002 Überlingen) where the risk was not spotted in time by controllers due to a calamitous chain of events. RAs were issued by both TCAS, one pilot followed its RA, the other followed ATC urgent descent instruction against RA. The BFU report noted "an automatic downlink, integrated in the TCAS equipment, carrying information about issued RA’s to the respective ATC units has not been activated yet" and has recommended to allow RA downlink to inform ATCOs:

To enhance the performance of ACAS
ICAO should initiate the development of down-linking RAs to ATC, using
such technologies as SSR Mode S and Automatic Dependent Surveillance -
Broadcast (ADS-B)

Studies on the feasibility and operational benefits of RA downlink have been conducted in the USA, in France (1994), and in Japan and lately from 2003 to 2007, with the FARADS project which concluded possible short-term downlink solutions would be useful:

With Mode-S RA reports, the controller will be aware about the RAs in
95% of the cases within 9 seconds of their occurrence. Thus, the
downlink of RA information was found to be sufficiently timely to
allow for a significant increase in the controller’s awareness of the
RA encounter.

In the course of this project, RADE-1 and RADE-2 experiments had taken place at Brétigny (France), and concluded in RA downlink benefiting to ATC, without creating a cognitive tunneling around the RA. it also indicated RA downlink could not be deployed without being included in existing ground-based safety nets.
ATC uses ground-based safety nets, which purposes are similar to aircraft ACAS and GPWS, e.g.

Short Term Conflict Alert (STCA): operational since around 2013. An alert in STCA is supposed to take place about 30s prior to a TCAS RA, however this is more complex, and the RA may happens before the STCA alert.

Area Proximity Warning (APW), Minimum Safe Altitude Warning (MSAW) or Approach Path Monitor (APM), all operational since around 2016

RADE-2 experiment showing labels with RA data on first line (source)
Eurocontrol project PASS studied the inclusion of ACAS RA into STCA, and more generally the mutual compatibility and interaction of both systems. In the final report (2010):

An optional real-time experiment with both controllers and pilots
in-the-loop was planned to complement the findings of Phase 1 and to
help define operational requirements supporting an overall operational
concept for joint  STCA/TCAS operations. This optional work was
however not launched due to lack of funding.

The work done concluded the interaction between STCA and ACAS should be considered by ANSP. A few early adopters (Czech Republic, Hungary, Japan, Luxembourg) are experimenting, to answer the question, yes the type of RA is displayed, and whether the aircraft is advised to climb or descend, as visible on the picture above. This is the current status, and obviously everybody is waiting for ICAO (representative of all countries) clarification on the topic.

Some ANSP's are experimenting with downlinking RA's to controller working positions via mode S. It will likely be implemented in some areas within a few years.

"RA takes precedence" only when the crew has been trained on the use of the system. I worked at a carrier that operated at least one airplane (mine) with RA capability but I was not legally able to take action on an RA that I received. The regs may have changed since but I recall this particular limitation becasuse it was just so stupid.

@acpilot: interesting :-) I can't answer for all jurisdictions, but at least in EU, ACAS II v7.1 are the only ACAS allowed, are mandatory for MTOW  5.7 t or 19 pax. Then EU-OPS 1.398 mandates SOPs to contain a requirement to use RA in flight, and "the commander or the pilot to whom conduct of the flight has been delegated must ensure that any corrective action indicated by the RA is initiated immediately".

It is technically feasible to alert ATC of the TCAS advisory via an automatic TCAS downlink:

"An automatic downlink for the TCAS – which would have alerted the controller that a TCAS advisory had been issued to the aircraft under
  his control, and notified him of the nature of that advisory – had not
  been deployed worldwide at the time of the accident." (source)

The above implies this system is available in some places in the world, but was not deployed worldwide in 2002 at the time of the Überlingen mid-air collision. I don't have information about the current status of worldwide deployment.


text:
Which aerobatic manuvers should I restrict myself to if I wear glasses?
Are there aerobatic maneuvers which I should never attempt if I wear glasses? Or, is there a specific G-loading which is not recommended for pilots who wear glasses?
I'm not into competitive aerobatics. I will not pull 6~7Gs or perform violent maneuvers like snap turns. However I do like to flip the world around just for the fun of it.

What sort of airplane are you flying?

ones where your glasses won't fall off

@J.Hougaard single-engine propeller monoplanes. Biplanes have too much rolling power (and obstruct my purpose of viewing the scenery upside-down).

Get glasses that won't fall off. You can get a strap that will attach to the glasses and go around your head keep them in place.

@rbp you hit the nail on the head in my opinion. That's the problem with this question.  This is nearly 100% opinion-based both as a question on stack exchange and as a pilot making decisions as the PIC.  Don't do it if you don't think you can do it safely.

Related on PPRuNe. Clearly opinions.

Depending on what type of headset you wear, you may not need to worry.  I wear a set of David Clark H10-13.4's, which have quite a bit of lateral clamping force on the sides of the head (unlike, for example, the Bose A20 and clones, which rest mostly on top of the head and have much less inward force around the ears).  I found that I was able to comfortably do all the basic maneuvers (loops, rolls, spins, hammerheads) as well as sustained inverted, negative 1G flight, without ever feeling like my glasses were in danger of leaving my face.
Without knowing more about your headset and the type of maneuvers you were doing (Mine were nowhere near what you'd call extreme) I can't be more specific, but you might be good to go just as you are.

If you're talking about what they call Gentleman's aerobatics - that is only controlled positive G maneuvers like aileron rolls, barrel rolls, Immelmanns, inside loops, split-S, Cuban Eights, etc. there shouldn't be a problem here as the loading will hold your glasses on your face.  For additional support you might consider a blackstrap for your eyewear, similar to what basketball players wear.
More extreme aerobatics involving negative G flight, tumbling, gyroscopic maneuvers, etc would not be recommended using free fitting corrective eyewear as they will probably come lose and can become a FOD hazard in the cockpit, in addition to losing your myopia correction.  Goggles with corrective lenses would be recommended here as well as contact lenses.


text:
How do airline pilots contact their company while in the air?
Sometimes airline pilots need to get in touch with their company while in the air, for example diagnosing cautionary messages or reporting a diversion. By what means is the communication channel established?

VHF / HF radio? If that's the case, who allocates the frequencies to each airline?
Satellite phone? (Are all aircrafts equipped with one in the cockpit?)
Something else?

I think I recall from reading other Q&As here that they do use satellite phone, but they might use other ways too though.

HF, VHF/UHF. A lot here about air-ground communications with different means. Overall ITU manages the allocation worldwide. US: FCC (member of the ITU conference) allocates frequencies to ARINC (Aeronautical Radio, Inc.), and ARINC deals with operators. See also: N4JRI's page. ACARS in addition of voice.

Related (not duplicate): What means do airline ground staff have to converse with pilots in flight?

While enroute the primary communication with company is via ACARS. (Aircraft Communications Addressing and Reporting System)
This is a digital datalink system and is kind of like sending text messages. On modern aircraft like a B777 it operates on HF, VHF, or SATCOM so there is coverage almost anywhere in the world.
If the aircraft is close to an airport with a company frequency, regular VHF radio communication can also be used.  Some airlines may even have company Peripheral stations (PAL) at certain locations so they can use a VHF radio to talk directly to Dispatch from hundreds of miles away.
In the USA, VHF company frequencies are allocated by Aviation Spectrum Resources, Inc. (ASRI) ASRI Ground Station Administration Service (AGSA)
SATCOM is also be used to make Satellite phone calls so we can talk directly to Dispatch or Maintenance from thousands of miles away. You can use pre-selected company numbers or manually enter any phone number.
Aircraft without SATCOM can use VHF or HF radios to contact ARINC (Aeronautical Radio, Incorporated) who can create a phone patch to any phone number.  Aeronautical Radio, Incorporated (ARINC)

Source: B777 Aircraft Operating Manual Volume 2 (2.23 Communications)

Great answer! If you would provide the source for your image, that would make it even better.

Thanks, I added the source. ( B777 Aircraft Operating Manual Volume 2 2.23 Communications)


text:
Do helicopters fly holding patterns under IFR?
If a helicopter is flying an IFR flight plan, and the airspace around the destination is busy, will controllers tell them to fly holding patterns just like planes do? Or will they hover at the holding fix? Or something else entirely?

Indeed helicopters can be put on holding during IFR.
ICAO DOC 8168 VOL II (PANS-OPS) has helicopter specifics in Section 4 Chapter 1 HOLDING CRITERIA:

1.3.2.2.2 Helicopter timing. The outbound timing should be:
a) one minute up to and including 1 830 m (6 000 ft); and
b) Category A fixed-wing aeroplane criteria above 1 830 m (6 000 ft).

Also Table II-4-1-2. Airspeeds for holding area construction specifies that for

Helicopters up to 1 830 m(6 000 ft) inclusive

airspeed is

185 km/h (100 kt)

under normal (non turbulent) conditions.
From the above, we can conclude that timing and airspeed equals flying the pattern and not hover the point.

In short, yes, helicopters fly procedure turns, hold, etc. just like airplanes.
Hovering at a fix is not practical, as they do not really have the nav capability to perform that, and it is contrary to normal ATC procedures.

It would also use more gas, not less. :)

AND have more flight risk.

I wonder what airships do as they can hover stationary without the fuel issue

Station keeping is an issue with an airship (they can't go sideways, and most of them backwards.  The instances I know of they fly the same as aircraft.  But they have ice issues, and usually travel VFR, at least the ones I have worked with.

A. A helicopter uses a LOT more fuel hovering than it does in forward flight.
B. ATC are used to seeing their radar blips moving, so having the helicopter keep on moving complies with radar's expectations.


text:
Is the Lift required for a steady-state climb and descent equal for the same absolute value of the kinematic climb angle?
Here is my confusion :
I found in my lecture script the following equation for the steady-state climb and descent (wings level)
L = m*g*cos(gamma) , with gamma being the kinematic climb angle (1)
(equation is simplified neglecting the Thrust angle of attack)

Now if an aircraft is flying at the same indicated airspeed (IAS) and the same absolute Vertical Speed (VS) once a climb (positive VS) and once a descent (negative VS), does it need the same lift in both situations ? (Neglect wind influence)
The kinematic climb angle should be equal in magnitude, but with flipped sign. According to this equation
gamma = asin(V_Vs./V_IAS); (2)

After plugging the values into (1) we get the same required Lift. Is therefore the angle of attack (AoA)  equal in the two scenarios ?
Also I assumed that:
Theta = AoA + gamma

Thank you for your help and clearing up my confusion :)

Hello Christian Mändle, and welcome aviation.stackexchange.com! Your question will be closed as a duplicate of Does lift equal weight in a climb?. The answer of Peter Kämpf addresses lift in descent as well. If there are

This question is clearly not the same as the claimed duplicates.  This question is about whether the effect is the same in climb and in descent.  This is a very reasonable answer because the usual explanation for the lift deficit in climb is that a component of thrust is offsetting the weight -- which is clearly not identically true in descent.
The short answer to the question is Yes.  $L=W, \cos(\theta)$ does not care the sign of theta.

Interestingly it is oppositely true in a descent:  the vertical component of drag is "offsetting" (some) of the weight


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Different Engine/Prop Sizes on Same Wing/Single Craft?
On a four propeller/engine wing, I’m wondering if theoretically at least, you could have a larger engine near the root of the wing with smaller engine(s) further out on the span..?.. If you could have the smaller propellers further out spin faster on smaller diameters and counter-rotate them... It's a pretty high wing so we can vary the prop diameter pretty well. The airframe design already has lots of vertical stabilization opportunities.
The engines would be diesel operating, generally at high torque at the 1500-2350 rpm range (further out on the wing engines would be rotating faster with high pitch at about 2200 rpm during cruise). On such a wing with different sized and powered engines would it be feasible to balance the torque-roll effects with computers (to calculate roll, yaw and pitch (relative to the NE Downframe?) angles on the fly), accelerometer sensors (arduino), spacing of the engines, construction of wings, vertical surfaces, tip tanks in the wings and air frame and rotating speed to rotating mass per diameter to feathered mass differentiation, etc.,?
Thanks so much,
Greg

While I'm not exactly sure what type of aircraft you're building, it sounds like it is quite a bit larger than a simple model aircraft. If this is true, I would strongly suggest not using an Arduino for any safety-related purpose. Arduino boards are meant for hobbyist use; they are not reliable enough for aircraft systems.

When you say " theoretical ", the answer is generally almost always yes. I don't see a real question in here though, mostly just musing about what might be possible. What is the purpose of this thought experiment?

The Bird Innovator was an American four-engined executive amphibious airplane modified from a Consolidated PBY Catalina by the Bird Corporation.
During the late 1960s Forrest Bird and his Bird Corporation developed a conversion for the PBY Catalina to improve performance.
The Catalina was modified by adding two 340 hp Lycoming GSO-480-B2D6 engines positioned outboard of the original Pratt & Whitney radials to increase range and performance.

No need to be so complicated, as long as you pair your engines, one on either side.
Putting a lighter engine nearer the wing tip may have been what Hugo Junkers had in mind with this aircraft, long before the computer age.
The V12 inboards swung 4 bladed props, the outer 6 cylinder engines had 2 bladed props.


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Does the “Torque-Induced Precession” of the two rotor blades cause the V-22 Osprey to tear itself apart
Concerning Bell Boeing V-22 Osprey Aircraft:
Does the “Torque-Induced Precession” of the two rotor -- 3-blade, propellers cause the V-22 Osprey to become unstable.
In the case of a helicopter the Earths gravity pulls on the aircraft hull up to the rotor assembly, rotor blade “dissymmetry of lift.”
Does the changing of the two gyroscopes, tiltrotor, cause “harmonic coupling” and or “mechanical resonance” - causing the aircraft to become unstable?

For this question, how does a tiltrotor differ from a non-tiltrotor two-rotor helicopter such as a Chinook?  How does it differ from a conventional twin-prop airplane?  (If it's about while the rotors are tilting, how does that differ from a Chinook or a twin changing pitch as a whole?)

"Does the “Torque-Induced Precession” of the two rotor blades cause the V-22 Osprey to tear itself apart?" Yes it does, all of the V-22 were destroyed because of that reason

Please note, “The second property is precession. Let’s follow the definition carefully. It is important to understand the effect of precession when we consider gyros. If an external force is applied to change the direction of the rotor spin axis -the gyro resists the change and instead moves the force at right angles to the direction originally applied, this is precession. The rule of precession states that the gyro will precess at 90 degrees to the applies force.
https://www.youtube.com/watch?v=JtniROSrovs

https://upload.wikimedia.org/wikipedia/commons/a/a0/V-22_concept.jpg
Please note,  If an external force is applied to change the direction of the rotor spin axis “Transmission Interconnect Shaft” the gyro resists the change and instead moves the force at right angles to the direction originally applied, this is precession.

How does a tiltrotor differ from a non-tiltrotor two-rotor helicopter such as a Chinook? How does it differ from a conventional twin-prop airplane? The answer is the  aircraft: angular momentum, structure and especially helicopters - stable center of gravity. Notice that the V-22 Osprey has two heavy engines relative to center of gravity.

Rotor blades are not rigidly attached to the rotor hub but via hinges which leave the blades freely move. There's no precession for a rotor rather a resonance between flapping and pitching movement. This is a well know phenomenon and that's why the blade control is 90° shifted in respect to the the blades.

The rotors of the V-22 Osprey turn in opposite directions.  This cancels and gyropscopic precession effects on the aircraft as a whole, eliminating the need for a tail rotor.
One may imagine a significant stress on each individual nacelle while yawing.  Ospreys are know for "gear box issues" and their production run may be over.


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Are there areas to avoid while flying formation with a larger aircraft?
Consider a fighter jet flying formation with an airliner-size aircraft, such as during an interception or mid-air refueling.
Are there areas which should be avoided by the fighter jet? I.e. are there certain positions where the downdraft or turbulence caused by the larger aircraft would send the small aircraft out of control?

This January 2017, a Bombardier Challenger rolled three times, losing 10,000 feet before it recovered when it crossed 1,000 feet under an Airbus A380.  It incurred structural damage during the recovery that resulted in total loss.

In case of air to air refueling, the 'receiver' (the aircraft being refueled) should be careful about not getting into the wake of the tanker, which is usually much larger and heavy. For example, the NATO air-to-air refueling manual, ATP-56 specifically states:

The  area  of  extreme  turbulence  directly  behind  and  slightly  to  the  right  of  the  tanker   should   be   avoided.   Blade   stall   and   uncontrolled   settling   may   be   encountered if this area is entered.

The receiving plane is usually flown below the wake turbulence of the tanker, in both the hose & drogue and flying boom methods. Usually, the receiver is directly behind and around 10' below the tanker. For example, see How do aircraft refueling in-air deal with wake turbulence?
In addition to turbulence, in case of close formation flying, there are other things to contend with, as explained in Advances in air to air refuelling:

In addition to the trailing air vortices the receiving aircraft must contend with downwash of airﬂow from the tanker. The affected local airﬂow typically exhibits a nose-down pitching moment on the receiver which, in turn, affects its speed and approach rate.

Not only the receiver, the tanker should also be careful about flow disturbances.

... mostly between large transport aircraft, the lead aircraft experiences a change in its airﬂow caused by the bow wave, or upwash, from the receiver. This increases the angle of attack around the tanker's tail which also results in a pitch-down motion.
Under-running the tanker is particularly dangerous since the downwash effect will dissipate, resulting in the receiver pitching up. At the same time a tanker on altitude hold may instigate a nose-down pitch due to a perceived climb due to the decrease in pressure in the surrounding air, potentially leading to collision.

The figure below shows some regions where disturbances would be encountered aft of a tanker.

Typical spanwise airﬂow disturbances from a trailing tanker wake.; image from Advances in air to air refuelling by Peter. R. Thomas et. al.
In addition, the trailing aircraft should also be careful about the jet exhaust.

It took me a looong time to get my head around that last graphic. I kept seeing the Downwash area to the left and the Backwash area to the right, as though that had been rotated 90° CCW. It wasn't until I started typing this that I realized that Downwash was, in fact, to the 'left', closer to the plane & vortex lines, and that Backwash was pointing away from the direction of flight. I knew that's how it should be, my eyes just refused to see it that way... Time for more coffee.

In higher altitude the wake turbulence moves downward behind the aircraft. So as the pilot of the lighter aircraft behind the bigger you should keep the same level at least.

photo by: Steve Morris, image source

I suppose the wake turbulence only becomes dangerous at some distance, when the downwash forms vortices. – Sanchises

No.
The wake turbulence starts forming at about 2/3 of the wing span (measured from body). It starts as a real narrow but powerful vortex. Then it raises its diameter losing power (slowly).

photo by: Josef  Willems, image source
As you can see near the aircraft the  wake turbulence has a sharp profile and is quite predictable. Therefore there is a safe area directly behind the tanker.
related post: How does an aircraft form wake turbulence?

Aerial refueling always takes place sightly underneath the tanker aircraft. I suppose the wake turbulence only becomes dangerous at some distance, when the downwash forms vortices.

Can you please add the source of the images and their copyright information? I have seen both pictures before, quite sure they are not free of copyright.

@DeltaLima at least the second one has been used at ASE before...

@DeltaLima thanks for the support!

@TimothyTruckle you're very welcome!

Dissimilar formation flying has numerous hazards, and not just with tankers
A classic example was the tragic accident in 1992 of a T-34 trainer getting too close to an F-14 while attempting to visually inspect a landing gear that had "unsafe" indications in the cockpit.  The T-34 is quite a bit smaller than an F-14.  Each aircraft will have a turbulent flow signature similar to the illustration aeroalias provided for the tanker in his answer (influenced obviously by airfoil shape and gross weight of the aircraft).  If you approach an aircraft that is not the same as yours, and are not familiar with it's wake areas, it is possible to make a fatal mistake.

14 Jan. 1992: ... a T-34C over Corpus Christi, TX. The Tomcat had an
  unsafe landing gear and the T-34 was checking it and later wrapped
  itself around the Tomcat's right wing. Both aviators in the T-34 were
  killed. The wreckage was left on the Tomcat's wing. The Tomcat land at
  the NAS there but had extensive damage.

Back when I was flying T-34's (some years before that accident) I did any number of gear checks on other T-34's and a couple of times T-44's at that same naval air station.  (The unsafe gear pattern was 2000' AGL).  What most likely caught the T-34 pilot in this case was using a similar approach (in terms of geometry) when joining up on an aircraft of considerably greater weight and a different turbulent signature, and then not having enough altitude to recover from the loss of control when the upset happened.
As the photographs illustrate, the disturbed air zones should inform how one makes an approach when closing with another aircraft.  There's a good reason aerial refueling is done at a high altitude.


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Are there any helicopters with ejection seats?
Are there any helicopters with ejection seats? If so, how is the clearance problem with the rotating blades solved?

First you blow off the blades, then the canopy and last the seat.

@PeterKämpf And hope you don't get an unwanted haircut. :)

As others above have stated, the Kamov KA-50 family is fitted with an ejection seat. First the blades are separated from the blade sleeves, which remain attached to the mast (rotor hub). a few fractions of a second later, the conventional upwards firing ejection seat is launched, using extraction rockets tied to cables to drag the seat clear of the airframe.
This cutaway image shows it to reasonable effect. '46' indicates where the blade separates from the sleeve. The little sketch of the smoke on the lower right blade shows the separation, and the '18' and '19' show the extraction rocket. KA-50 Cutaway image
Digital Combat Simulations's (DCS) line of computer simulations models the ka-50, including the ejection seat. I'm sure there's video from the game online somewhere.

I guess we don't have to worry about one of these landing itself in a cornfield after pilot ejection (as happened once with an F-106)...

Many years ago I read about the Kamov-50 helicopter family* being the first helicopters equipped with ejection seats. And Wikipedia verifies my memory:

It is the world's first operational helicopter with a rescue ejection system, which allows the pilot to escape at all altitudes and speeds.

In the same article for the same helicopter it is mentioned that

Before the rocket in the ejection seat deploys, the rotor blades are blown away by explosive charges in the rotor disc and the canopy is jettisoned.

Also, if you want to count experimental helicopters in, Sikorsky S-72 RSRA may be another one. Here is an interesting video (facebook link), demonstrating the blades being blown away, and ejection of the pilots.

As a side note, in the article I read in an aviation magazine many years ago, it mentioned a synchronization of the seat with the blade rotation so that the seat passes between the blades. It was based on a working principle similar to the middle machine gun that some WW2 German airplanes were equipped with that fired the projectiles between the propeller blades.
Judging from the fact that I read it somewhere in the late 90s where the project was still new and developing, as PlasmaHH pointed out, that was probably more a thought rather than a fact.
* Kamov 50 and its variants/successors like Ka52

The Ka-52 (the Ka-50s successor) is equipped with ejection seats too (Swesda K-37). I can't imagine that the "shoot through the blades" was more than an idea, given the necessary speed for that (which results in a probably very unhealthy acceleration). Similarly there was once the idea evaluated to reinforce the seat so that when it collides with the blades will protect the pilot(s).

To do some math on the blade sync, getting a 1m high sitting human through blades going past at 300 RPM, twice per rotation, 10 times a second, requires a speed of only 10m/s. An acceleration calculator says a 5g acceleration would give that speed with only 1m travelled. I think there's enough margin that this could work at least in theory.

Does the downvoter care to explain? Or I'm asking for too much?

That comma in the first Wikipedia quote feels dodgy. Does a rescue ejection system in a helicopter necessarily allow escape at all altitudes and speeds? Or is this the first that had a rescue ejection system which had the property of allowing such escape? Or what? IIRC at least in fixed-wing aircraft, some (particularly earlier generation) ejection seats required a minimum altitude to be safe (for whatever definition of "safe" is in use) to use as the parachute needed time to deploy and slow the pilot to a safe landing speed, while keeping initial accelerations survivable.

What you need is an old-style crop-dusting helicopter. No ejection seat required - there were no doors at all, to maximize visibility sideways and down. You checked your seatbelt very carefully before takeoff! (I've flown as a passenger in one of those, doing some aerial photography - scary, when you are hanging out of the door to get a better field of view and the pilot banks 30 degrees to make a tight turn!

@MichaelKjörling Looking at the video link, it appears to be launched from the ground in the test, and the parachutes look like they engage.  I am in no way qualified to make a judgement, but it looks like there is some merit to "all altitudes" just based on that.

If I recall correctly, those WW2 synchronizers were often less than 100% reliable (for many different reasons).  Not necessarily a design you'd want to trust in an ejection system.  Even if the necessary ejection velocity is technically attainable.

@aroth: Compared to all the other ways an ejection seat can kill you, imperfect synchronisation might not raise the lethality of the system that much.

Ah the lovely internet and its volatility. Thanks @RodrigodeAzevedo, I found the same video on FB and I replaced the broken link. Until the next time...

Image source
The Sikorsky S-72 Rotor Systems Research Aircraft had ejection seats, as shown on this video. Like the Ka-50, it blew the rotor blades off first before ejecting.

This article mentions a crash of a Mi-28 in which one pilot died:

The Mi-28 was supposedly designed with an ejection seat system that fires its crew out the side and downward.

Out the side?!  Surely that wasn't fully thought through... how would it not instantly break your neck?

@J... Yeah man, doesn't sound right, no wonder if it was never implemented.

Firing down isn't particularly good for low-level ejection. :(

The Mi-28 doesn't actually have the ejection seats, the Mikoyan constructor bureau follows a different approach: they've created energy absorbing seats and landing gear to protect the pilot in a crash.
Kamov does install ejection seats in their helicopters, a.o. in the Ка-52. It blows off the rotor blades and shatters the upper glass windows to make the way clear for pilot ejection. One can see the white cord with explosive on the upper window of Ka-52 on the pictures (there is a white zigzag-like cord on the bottom of the pilot window):
Testing of the part of the system

What is "..amortization for seats and chassis"?

I would imagine he meant armor.

@Davidw Ah, that makes sense.

Nope, it is not about armor. Check out chassis of Mi-28 in the message above, it has pretty developed amortization part : hydraulic\pneumatic cylinder is attached to each of front chassis. In publications it is stated that combined with others impact  absorbing elements of construction it should reduce the impact to "physiologically acceptable" level for speeds up to 12 meters per second. Additional concerns are rise against fire safety for fuel tanks (no such separated system develped in russian helicopters yet).

Mi-28 is pretty good at fire safety, though, no fire happened during several pretty sever crashes: one crash
another

Heh, just realized that I use wrong term chassis for landing gear. I actually meant front landing wheels of Mi-28, not the armored capsule.

The Mi-28 has no ejection seat, but the crew is equipped with parachutes and there is an inflatable slide to help them. Take a look at this.


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What is the requirement to become a flight attendant?
We know that flight attendants, on top of serving passengers (which is what they do 99% of the time), are there to ensure the safety of passengers on board. For example, they are trained to operate emergency equipment in the event it is necessary.
What is the legal requirement to become a flight attendant? I am not asking how one can get hired (which may involve speaking different languages, being friendly to customers etc.), but how one can serve as a flight attendant on a commercial flight recognized by aviation authority.

Flight attendants get a bad rap, but they are important required crew members on passenger flights. Flight attendants are trained for various emergencies. I've actually seen them doing emergency water training at a university pool where they were training on the proper procedure for entering a life boat and pulling people into the lifeboat. It looked like some serious training to me. They each took turns being in the water and in the life boat.

There is a (relatively) new requirement for Flight Attendants (on aircraft with more than 20 passenger seats) to be certified as per the FAA. This was done in the "VISION 100—CENTURY OF AVIATION
REAUTHORIZATION ACT", in section 814 if you want to read it.
The applicable CFR is 49 U.S. Code § 44728 - Flight attendant certification. In summary it says:

You have to have a certificate of demonstrated proficiency, this looks like a pilots license.
Demonstrated compliance with the air carrier approved training program.
Proficiency to read, speak, and write English
Comprehend written material in English
Ability to write incident reports and log entries/statements
Carry out written and oral instructions

Now, the actual physical requirements are up to the airline, as long as the individual can complete the training program, they are physically allowed to become a flight attendant (physical appearance requirements aside). As for "legally" allowed, they need to be able to pass the same TSA background checks that pilots are subject to. So legally, almost all airlines require:

U.S. Citizen or appropriate work visa
Be able to enter/exit the United States without any issues (no visa problems with other countries, although for regionals this may not apply)
Be able to pass a TSA/FBI background check
No DUI or felony convictions within the past 10 years

There isn't a lot more "legal" to it than that, they don't have to demonstrate the use of safety equipment if the airline's training program doesn't require it.

These requirements can vary significantly depending on different airlines in different countries. Since you added faa-regulations, I'll assume you are only interested in US based airlines.
According to TheTravelAcademy, the requirements are:

PHYSICAL

NO VISIBLE TATTOOS
4’11″-6’4” tall- These numbers are based on various airline requirements.
Excellent overall health
Five senses- hearing/sight/touch/smell/taste
Weight and height proportionate
An overall pleasing and well groomed appearance
Vision that is correctable by contacts or glasses
No facial piercing or gauges- 1 earring per ear (lobe only)
At least 19 years old*

EDUCATION/LANGUAGE/EXPERIENCE

Minimum high school diploma or GED- Any additional education is a plus!
Customer service experience – At least 2 years is preferable and in some cases required.
Fluent in English (reading, writing, listening, and speaking)- Bilingual people must read, write, understand, and speak English and
  an additional language fluently. Second language proficiency will be
  tested.

CITIZENSHIP/IDENTIFICATION/BACKGROUND

A United States citizen- When applying to a US-based airline, applicants must have full legal ability to work in the US and be able
  to exit and reenter US without incident.
Identification- This includes a valid passport, social security card, and/or government issued picture ID
A clean background- Flight attendants may not have any felony convictions, DUI convictions within the last 10 years. They must also
  pass an FBI background check.

FlightAttendantCareer also listed similar requirements in their FAQs.

Why would the FAA care about tattoos (in capital letters, even)?

@HenningMakholm These requirements (other than the citizenship/background/language) are a composite of different airline requirements, not regulatory. Those other requirements are put together by the "The Travel Academy".


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Why are the controls of transport category aircraft much heavier than light aircraft?
Anyone who has transitioned from light GA aircraft to large turbofan airliners will notice the difference: the control column on a large aircraft is much heavier. While still within limits of a healthy adult, it takes a considerably larger force to move the controls.
Why is this? Is this intentional, or just a by-product of the early days mechanical linkages? Airbus has shown that even on airliners, the controls can be made as light as simple joystick. So why can't the controls of a Boeing 747 feel the same as a Cessna 172?

The control forces are there by design.  All airliners have artificial force feel from springs, dampers etc. A B747 flies through the skies a lot faster than a Cessna 172 and bad things could happen if it were easy to deflect the flying controls at M 0.85.
The B737 elevator is a good case in point. It has q-feel (artificial feel) provided by a spring device, and the spring gradients stiffen up as a function of airspeed.  At higher speeds it is quite hard to deflect the column by pushing/pulling and much easier to use the trim button to control the pitch. For Level D simulator control forces we measured the forces necessary to deflect the column fully when trimmed in mid position, which turned out to be a serious body-building exercise:

About 200 N at zero airspeed
About 500 N at 400 knots 20,000 ft. (Measurement taken on ground with a pressure hose attached to the pitot tube, and aircraft on jacks with the landing gear retracted).

By contrast, helicopter flying controls are light because helicopters are inherently unstable and a lot of stick action is necessary, particularly in the hover. In most helicopters the feel forces can be switched off completely with the trim button, resulting in a very light stick, perfectly OK with the relatively docile response of the helicopter.
But a 747 is a stable airframe. A fast aircraft with large control deflections results in high accelerations, which can result in Pilot Induced Oscillations or structural damage to the airframe.

The two measurements of B737 control column, digitised from the original plots and on the same scale: column position on the x-axis, force on the y-axis. By design, force required to deflect the column fully from about mid position is about 40 lbf at zero knots, 100 lbf at 400 knots. This is from an artificial q-feel system and the feel is implemented by design. The vertical distance between the sides of the curve is caused by friction.

At 0 knots:

Notice that around mid the spring gradient is very steep: the breakout, to indicate neutral position.
Then between 0 and 5 inch a spring gradient of 4 lbf/inch.
Beyond 5 inch a relatively flat gradient, to make it not too hard to deflect the column fully.

At 400 knots: it is made hard to fully deflect the column. 100 lbf pull from a seated position is nothing to sneeze at.

Unit conversions: 200 N is about 45 lbs, 500 N is about 110 lbs.

I'm curious is that 400 knots indicated or true airspeed? At 400 KIAS a B737 would have disintegrated...

@kevin The measurements were taken on ground, with a measurement set with a position and force transducer attached to the column. A pressure hose was attached to the pitot tube: 400 KIAS. And then an untinterrupted, continuous, slow sweep of the control, starting from neutral, push slowly until in the fwd stop, release slowly then pull into aft stop, then relese slowly again. The plot is of control force vs position and shows the spring gradients, friction, travel limits, cable stiffness.

Adding to what @TannerSwett wrote above, for those of us more used to SI units but who don't necessarily have an intuitive grasp of newtons, it's often close enough to just say that 10 N = 1 kgf, so divide forces in newtons by 10 and you get the equivalent force in kgf. I'll be the first to admit that this is not a perfect equivalence, but unless you are writing a scientific or detailed engineering paper, it's good enough for most purposes. The exact value is 9.80665 newton per kgf.

@TannerSwett correction: pounds of force (lbf), not pounds (lb).  The former is force, the latter is mass.

Most people don't realize that most airliners fly very close to but not exceeding Mach 1

@slebetman: Most people on Aviation.SE do.

Is the 737's artificial feel still active in manual-reversion flight (where the aerodynamic loads imposed by the airflow over the control surfaces on the deflected aileron and elevator servo tabs cause the controls to stiffen up all on their own with increasing airspeed), where the artificial feel would (at best) be redundant, and (at worst) impair aircraft controllability by excessively limiting flight-control deflection?

@Sean with all hydraulic power off, the artificial q-feel forces are zero. There is increased friction and direct feedback of the aeroforces over the control surfaces.

Beyond 5 inches at 0 kt, flatter gradient would cause easier movement rather than harder movement according to graph. You’re getting more deflection with the same amount of force in increase. This graphs are kind of weird though. Normally force gradient should increase close to full deflection (steeper gradient) to avoid over stress the aircraft caused by lighter control column forces.